Satellites and satellite fleet implementation methods and apparatus

ABSTRACT

A method for implementing a satellite fleet includes launching a group of satellites within a launch vehicle. In an embodiment, the satellites are structurally connected together through satellite outer load paths. After separation from the launch vehicle, nodal separation between the satellites is established by allowing one or more of the satellites to drift at one or more orbits having apogee altitudes below an operational orbit apogee altitude. A satellite is maintained in an ecliptic normal attitude during its operational life, in an embodiment. The satellite&#39;s orbit is efficiently maintained by a combination of axial, radial, and canted thrusters, in an embodiment. Satellite embodiments include a payload subsystem, a bus subsystem, an outer load path support structure, antenna assembly orientation mechanisms, an attitude control subsystem adapted to maintain the satellite in the ecliptic normal attitude, and an orbit maintenance/propulsion subsystem adapted to maintain the satellite&#39;s orbit.

This application is a divisional of application Ser. No. 11/693,645,filed Mar. 29, 2007, status pending.

TECHNICAL FIELD

Embodiments described herein generally relate to satellites and tomethods and apparatus for implementing satellite fleets.

BACKGROUND

Satellite systems have been used extensively to provide a variety ofsatellite-based services and functions. For example, current satellitesystems provide communications services, broadcast and multi-castservices, earth imaging, radar, weather system monitoring, andastronomical observations, among other things.

A satellite system may consist of a single satellite or a plurality ofsatellites (e.g., a “satellite fleet”) whose functionality are combinedin some manner. Satellites may travel in geostationary orbits (GSO) ornon-geostationary orbits (NGSO), and are referred to as GSO satellitesor NGSO satellites, respectively. A satellite of a single satellitesystem typically is a GSO satellite, so that the satellite may providecontinuous service over a distinct coverage area on the surface of theearth. Satellite fleets may include GSO and/or NGSO satellites. In orderto provide continuous coverage, the orbital planes of the satelliteswithin an NGSO satellite fleet typically are separated by some nodalseparation angle.

To deploy a satellite fleet, satellites are placed into orbit usinglaunch vehicles. In some cases, a single satellite is carried into orbitby a single launch vehicle. Because launch vehicles and launchoperations are expensive, implementing a satellite fleet usingsingle-satellite launches may be uneconomical, in some cases.Accordingly, some satellite fleets have been implemented by launchingmultiple satellites within a single launch vehicle. In amultiple-satellite launch, the satellites are structurally connectedtogether using a special adapter or dispenser. Although some economiesare gained through multiple-satellite launches, the special adapters ordispensers add launch mass and cost. Additionally, once the satellitesare released from the launch vehicle (e.g., the payload fairing), theyoperate in the same plane or at least some of the satellites performbooster rocket burn operations to establish appropriate nodal separationbetween the satellites' orbital planes. To have the fuel capacity toperform the booster rocket burn operations, satellite fuel tanks aremade to accommodate the extra fuel, which also adds extra launch mass.

NGSO operation may have the benefit of improved coverage for higherlatitudes. However, NGSO operation suffers from a wider range solar betaangles (i.e., the angle between a satellite's primary axis and thedirection of the sun). Because it is desirable to orient a satellite'ssolar panels perpendicular to the direction of the sun whilesimultaneously pointing the satellite's downlink antenna toward theintended coverage area, NGSO satellites include relatively complexsteering mechanism and multiple gimbals to achieve solar panelorientation and downlink antenna steering.

It is desirable to provide methods and apparatus for economically andreliably implementing a satellite fleet to provide satellite-basedservices. In addition, it is desirable to provide satellites capable ofproviding the satellite-based services, which have reduced mass and/orcomplexity when compared with traditional satellites. Other desirablefeatures and characteristics of embodiments of the inventive subjectmatter will become apparent from the subsequent detailed description andthe appended claims, taken in conjunction with the accompanying drawingsand the foregoing technical field and background.

BRIEF DESCRIPTION OF THE DRAWINGS

Various embodiments will hereinafter be described in conjunction withthe following drawing figures, wherein like numerals denote likeelements, and

FIG. 1 illustrates a portion of a direct broadcast satellite system, inaccordance with an example embodiment of the inventive subject matter;

FIG. 2 illustrates a ground track, in accordance with an exampleembodiment;

FIG. 3 illustrates a coverage area, in accordance with an exampleembodiment;

FIG. 4 illustrates a configuration of orbits for multiple satellitesfrom the perspective of space, in accordance with an example embodiment;

FIG. 5 illustrates a cross-sectional, side view of a satellite, inaccordance with an example embodiment;

FIG. 6 illustrates a perspective, cut-away view of a satellite, inaccordance with an example embodiment;

FIGS. 7-9 illustrate perspective top views and cut-away side views of asatellite with a antenna assembly in first, second, and third positions,in accordance with an example embodiment;

FIG. 10 illustrates a perspective view of a satellite, in accordancewith an example embodiment;

FIG. 11 illustrates a top view of a satellite, in accordance with anexample embodiment;

FIG. 12 illustrates a flowchart of a method for implementing a satellitefleet, in accordance with an example embodiment;

FIG. 13 illustrates a cross-sectional, side view of a configuration ofmultiple satellites stacked within a payload fairing of a launchvehicle, in accordance with an example embodiment;

FIG. 14 illustrates a perspective view of a configuration of multiplesatellites stacked within a cut-away payload fairing of a launchvehicle, in accordance with an example embodiment;

FIG. 15 illustrates a simplified diagram of a propulsion subsystem, inaccordance with an example embodiment;

FIG. 16 illustrates a simplified bottom view of a satellite,illustrating a rocket booster, axial thrusters, and radial thrusters, inaccordance with an example embodiment;

FIG. 17 illustrates a simplified side view of a satellite, illustratingcanted thrusters, in accordance with an example embodiment; and

FIG. 18 illustrates a simplified block diagram of various satellitesubsystems, in accordance with an example embodiment.

DETAILED DESCRIPTION

The following detailed description is merely exemplary in nature and isnot intended to limit the described embodiments or the application anduses of the described embodiments. Furthermore, there is no intention tobe bound by any expressed or implied theory presented in the precedingtechnical field, background, brief summary or the following detaileddescription. An example satellite fleet is discussed, below, whichincludes six NGSO satellites in Molniya orbits, which provide directbroadcast satellite (DBS) services. It is to be understood that thedescription of the example satellite fleet is not intended to limit thescope of the inventive subject matter to such a fleet. Instead,embodiments of the inventive subject matter may be used in GSOsatellites, NGSO satellites, satellites traveling in orbits that aredifferent from Molniya orbits, and/or satellite fleets having anywherefrom two to numerous satellites, and/or satellites that provide othertypes of services or perform other types of functions. Accordingly, thescope of the inventive subject matter is intended to include a varietyof satellite and satellite fleet embodiments.

Embodiments of the inventive subject matter include satellites andmethods of implementing satellite fleets. Embodiments of the inventivesubject matter may be implemented in any of a variety of types ofsatellites (e.g., GSO satellites and/or NGSO satellites) and in any of avariety of satellite fleets. An example satellite fleet is discussed inconjunction with FIGS. 1-4. In particular, an example satellite fleetmay form a portion of a direct broadcast satellite (DBS) system. Theplurality of satellites within the fleet may transmit (e.g., broadcast)television signals toward a target geographical area. In an embodiment,each satellite may transmit television signals within the C-band (e.g.,within a frequency range between 3.7 and 4.2 GHz), although satellitesmay transmit signals within other bands, in other embodiments. Theplurality of satellites are in Molniya orbits, in an embodiment, whosenodes are substantially equally separated around the earth. Thesatellites are phased in their orbit paths so that they track along acommon ground track.

FIG. 1 illustrates a portion of a DBS system 100, in accordance with anexample embodiment of the inventive subject matter. DBS system 100includes a plurality of satellites, such as satellite 102, within theuser beams, at least one uplink hub 104, and a plurality of userequipment (UE) systems, where each UE system includes a UE antenna 106.Although FIG. 1 shows a single satellite 102, DBS system 100 may includea plurality of satellites. In an embodiment, DBS system 100 includes sixsatellites, although a system that implements an embodiment of theinventive subject matter may include more or fewer satellites. Inaddition, although one uplink hub 104 and three UE antennas 106 areillustrated, a system may include more or fewer (including zero) uplinkhubs and/or UE antennas 106. Systems that implement embodiments of theinventive subject matter may include other types of equipment, as well.

Uplink hub 104 may include a control facility and one or more uplinkantennas, in an embodiment. Uplink hub 104 is adapted to transmit uplinksignals 110 toward satellite 102. In an embodiment in which DBS systemincludes a plurality of satellites, uplink hub 104 may transmit uplinksignals 110 toward each satellite when the satellite is in view ofuplink hub 104. Uplink signals may include, for example, satellitecontrol information and/or content, among other things. In anembodiment, content may include uplink television signals, which mayinclude multiple audio and video content streams.

Satellite 102 is adapted to receive uplink signals 110 from uplink hub104. Satellite 102 further is adapted to transmit downlink signalstoward the surface of the earth. Uplink and downlink signals may includecontent, such as television signals, for example. In an embodiment,satellite 102 transmits downlink signals in at least one directed beam112. Although FIG. 1 illustrates a single beam 112, satellite 102 maytransmit downlink signals in multiple directed beams, which may or maynot overlap each other. Signals transmitted within a directed beam 112may intersect the surface of the earth within a region 114. A region 114may define a geographical area within which the signal strength and/orsignal-to-noise ratio (SNR) of the transmitted downlink signals isconsidered to be above a threshold. A region 114 may be substantiallycircular, elliptical, or otherwise shaped, in various embodiments.

In various embodiments, satellite 102 may multi-cast and/or broadcastdownlink signals toward a population of UE systems that are locatedproximate to a surface of the earth. In a particular embodiment,satellite 102 is adapted to receive uplink television signals and tobroadcast downlink television signals, which UE antennas 106 and systemsare adapted to receive. Satellite 102 transmits downlink televisionsignals within a frequency range between 3.7 and 4.2 Gigahertz (GHz), orwithin the C-band, in an embodiment. In other embodiments, satellite 102may transmit downlink signals within other bands and/or within otherfrequency ranges.

As will be described in more detail in conjunction with FIG. 2,embodiments of a system may include a plurality of satellites (e.g., a“satellite fleet”), which follow highly elliptical orbit (HEO) paths(e.g., Molniya orbit paths). In other embodiments, a system may includea plurality of satellites that follow other NGSO paths, and/or GSOpaths.

As used herein, the term “Molniya orbit” means an orbit having aninclination of about 63.4 degrees, for which there is substantially noprecession of the line of apsides. A Molniya orbital period of abouttwelve hours is synchronized with that of the earth's rotation, with tworevolutions per day. Thus, two apogees with longitudes 180 degrees apartremain substantially fixed relative to the earth. In variousembodiments, the orbit has an eccentricity within a range of about 0.71to 0.75 (e.g., about 0.729 in a particular embodiment). In addition, inan embodiment, the orbit has an apogee altitude in a range of about39,000 to 40,000 kilometers (km) (e.g., about 39,547 km in a particularembodiment), and a perigee altitude in a range of about 260 to 1325 km(e.g., about 820 km in a particular embodiment). The term “Molniya”orbit, as used herein, is not meant to limit the scope of the inventivesubject matter to any other external definition of the term.

At an orbit inclination angle of 63.4, the absidial line (i.e., the linecontaining the semi-major axis of the ellipse) remains oriented in aboutthe same direction in inertial space. Accordingly, this angle ofinclination substantially prevents “absidial drift,” or precession ofthe line of apsides around the orbit. In various embodiments, the orbitsmay have apogees in either the Northern or Southern hemispheres.

In an embodiment, an orbit period is about one-half of the sidereal day,in an embodiment, or about 12 hours. Accordingly, the orbit issubsynchronous with the rotation of the earth. Due to thesubsynchronicity of the orbit, a ground track may include two zeniths inthe Northern hemisphere (or two nadirs in the Southern hemisphere), eachof which is associated with an apogee of the satellite's orbit. Theground track zeniths are substantially stationary, meaning that each oneis located at a fixed latitude and longitude. Accordingly, each day, asatellite reaches a first apogee over an intersection between a groundtrack first latitude and a first longitude, and a second apogee over anintersection between a ground track second latitude and a secondlongitude.

In an embodiment, the satellites within the fleet have orbit parametersthat cause the satellites to present themselves over substantially thesame geographical areas during their orbit rotations. In other words,the satellites may follow orbit paths associated with a substantiallyfixed ground track. This ground track is referred to herein as a“common” ground track, because all satellites within the fleet arephased in their orbits so that they follow the same ground track. Inanother embodiment, the satellites within a fleet have orbit parametersthat do not result in the satellites following orbit paths associatedwith a common ground track. In still another embodiment, the satelliteswithin a fleet may include GSO satellites, which remain substantiallyfixed with respect to the surface of the earth.

FIG. 2 illustrates a ground track 200, in accordance with an exampleembodiment. Ground track 200 represents a substantially fixed path onthe surface of the earth, over which each satellite of a satellite fleettravels during its orbit. As discussed above, ground track 200 mayinclude two zeniths 204, 206, each of which corresponds to an apogee ofa satellite's orbit. In an embodiment, each zenith 204, 206 is locatedat a fixed latitude 210 and a fixed longitude 212, 214. In anembodiment, latitude 210 is about 63.4 degrees North latitude for bothzeniths 204, 206. A first zenith 204 of ground track 200 is located at afirst longitude 212 (e.g., a longitude bisecting Europe or a longitudeof about 10 degrees east being preferred in an example embodiment). Asecond longitude 214 may correspond to a second zenith 206 that is about180 degrees separated from the first longitude 212, in an embodiment.Although the example embodiment of FIG. 2 describes preferredlongitudes, any of a wide range of longitudes may be chosen in which thecommunications service is supported.

A coverage area may correspond to a geographic area on the surface ofthe earth within which satellite services are provided. FIG. 3illustrates a coverage area, in accordance with an example embodiment. Acoverage area may be represented by one or more contours. For example,FIG. 3 illustrates six contours 301, 302, 303, 304, 305, 306. Eachcontour may correspond to a predefined signal power and/orsignal-to-noise ratio, where a signal-to-noise ratio at an inner contour(e.g., contour 301) may be higher than a signal-to-noise ratio at anouter contour (e.g., contour 306). The flatness of the pattern may beaffected by various antenna characteristics. For example, a satellitemay include a shaped or distorted parabolic reflector (e.g., reflector542, FIG. 5), which may project a downlink beam (e.g., beam 112, FIG.1). Desirably, the gain is optimized over the desired coverage area. Aswill be described in more detail later, a satellite may additionallyinclude one or more mechanisms for aiming the beam at the desiredcoverage area. The contours 301-306 of FIG. 3 illustrate the gain of theantenna for transmission, as viewed by a satellite.

Within a coverage area, a DBS system may provide a downlink capacitywithin a range of about 800-900 Megabits per second (Mbps), with acapacity of about 864 Mbps in a particular embodiment. This correspondsto 36 Mbps for each of 24 RF channels, in an embodiment. Each 36 Mbpschannel may contain a plurality of ordinary and/or high-definitiondigitally compressed television signals (e.g., 18 ordinary and/or 7high-definition digitally compressed television signals). When thesystem services two coverage areas (e.g., areas proximate to two apogees204, 206, FIG. 2), the system may provide a total downlink capacity thatis about twice the capacity of a single coverage area, or within a rangeof about 1.6 to 1.8 Gigabits per second (Gbps), with a total capacity ofabout 1.728 Gbps in a particular embodiment. A system may providedownlink capacities larger or smaller than the above given ranges, inother embodiments.

FIG. 4 illustrates a configuration of orbits for multiple satellitesfrom the perspective of space, in accordance with an example embodiment.Circle 400 represents the earth as seen from a point far above the NorthPole. In a particular embodiment, a satellite fleet includes sixsatellites, and the ellipses in FIG. 4 represent orbit paths 401, 402,403, 404, 405, 406 for the six satellites, where the orbit paths 401-406have about a 60 degree nodal separation. The orbital planes for thesatellites may be considered to have “widely separated nodes,” where theterm “widely separated nodes,” as used herein, means a nodal separationgreater than about 10 degrees. In various embodiments, satellites of asatellite fleet may have nodal separations in a range of about 10degrees to about 180 degrees. Orbit paths 401-406 correspond to Molniyaorbits, in an embodiment. Apogees 411, 412, 413, 414, 415, 416 andperigees 421, 422, 423, 424, 425, 426 are indicated with diamonds andcrosses, respectively, for each orbit path 401-406. In otherembodiments, a satellite fleet may include more or fewer than sixsatellites, the orbit paths may have something other than about a 60degree nodal separation, and/or the orbit paths may correspond to otherthan Molniya orbits.

In order to provide satellite services in accordance with embodimentspreviously discussed, the system infrastructure first is established.This infrastructure includes a fleet of satellites, in an embodiment. Inother embodiments, a system may include as few as one satellite. As willbe described in detail below, satellites of various embodiments includevarious design features that may provide certain advantages in the areasof satellite fleet implementation and operations, among other things.

FIG. 5 illustrates a cross-sectional, side view of a satellite 500, inaccordance with an example embodiment. Generally, satellite 500 includesan outer load path support structure 502, a solar energy collectionsubsystem, a bus subsystem, and a communications payload subsystem.Outer load path support structure 502 includes a rigid and substantiallycylindrical structure that defines a primary axis 504 of satellite 500.The term “substantially cylindrical” means, in various embodiments, thatouter load path support structure may have a substantially circularcross-section, or a cross-section having a different geometric shapethat enables the outer load path support structure to substantiallysurround the bus subsystem and the communications payload subsystem.

In an embodiment, the attitude of satellite 500 is maintained so thatprimary axis 504 remains substantially perpendicular to the eclipticplane (i.e., the geometric plane containing the mean orbit of the eartharound the Sun). This attitude is referred to herein as an eclipticnormal attitude. Maintenance of a satellite in an ecliptic normalattitude may provide several advantages over traditional systems, whichtypically maintain their satellites in an orbit normal attitude (i.e.,an attitude with a primary axis perpendicular to the satellite's orbitplane), an equatorial normal attitude (i.e., an attitude with a primaryaxis parallel to the equatorial plane) or in a sun nadir steeredconfiguration.

As will be described in more detail later, outer load path supportstructure 502 is adapted to provide structural support for stackingmultiple satellites in the payload fairing (e.g., payload fairing 1304,FIG. 13) of a booster rocket. In an embodiment, outer load path supportstructure 502 has a diameter (or width) in a range of about 2 m to about5 m, with a diameter of about 3.5 m in a particular embodiment. Outerload path support structure 502 includes structural members (notillustrated) to define its shape, an exterior surface 506, an interiorcavity 508, and an earth opening 510, in an embodiment. The exteriorsurface 506 may be substantially solid, or may include apertures. Theearth opening 510 exposes a reflector 542 of the communications payloadsubsystem, and during operations, the earth opening 510 is directedtoward the ecliptic while the exposed antenna is pointed towards earth.Outer load path support structure 502 also may have a space-orientedopening at an opposite end from earth opening 510, in an embodiment. Inan alternate embodiment, satellite 500 may include an end cap structure512 at the opposite end.

The solar energy collection subsystem includes one or more solar cellassemblies 514, which convert solar energy into electrical energy. Inaddition, the solar energy collection subsystem includes at least onebattery (not illustrated) that receives and stores the electrical energyproduced by solar cell assemblies 514. This energy is consumed, duringoperations, by various elements of satellite 500. In an embodiment,solar cell assemblies 514 may generate a substantial amount of power(e.g., about 2 kilowatts) for consumption by the satellite elements.

In an embodiment, solar panels of the solar cell assemblies 514 areattached to and/or define a portion of the exterior surface 506 of outerload path support structure 502. Because the primary axis 504 ofsatellite 500 is maintained substantially perpendicular to the eclipticplane, in an embodiment, the sun impinges directly on exterior surface506 throughout the satellite's entire orbit, except when the satelliteis in the shadow of the earth or moon. Additionally, the interior cavity508 is minimally exposed to the sun throughout the satellite's orbit.Accordingly, within the interior cavity 508 of outer load path supportstructure 502, significant thermal variations due to solar heating arenot likely to occur. A satellite design, in accordance with anembodiment, may result in a substantially consistent thermal load withinthe interior cavity 508 of the satellite, when compared with satellitesthat maintain orbit normal, equatorial normal or sun nadir steeredattitudes. In other words, maintenance of satellite 500 in an eclipticnormal attitude may result in a thermal boundary that is approximatelyconstant throughout each orbit and throughout the operational life ofsatellite 500. This may enable a simplified thermal load controlsubsystem to be implemented on-board satellite 500, when compared withtraditional satellites that experience widely varying thermal loads. Inaddition, traditional satellites include various mechanisms (e.g.,gimbals) designed to dynamically adjust the orientation of the solarpanels, with respect to the satellite body, in order to keep the solarpanels facing the sun. Traditional satellites also may orient the entiresatellite body to prevent excess thermal load. Because solar panels ofthe solar cell assemblies 514 inherently face the sun at virtually alltimes during system operations, and the earth opening 510 andspace-oriented opening (or end cap structure 512) of the satellite neverhave solar impingement, such mechanisms and body steering may beexcluded according to embodiments of the inventive subject matter.

The bus subsystem is located substantially within interior cavity 508generally toward end cap structure 512 (or a space-oriented opening),and is rigidly coupled to the outer load path support structure 502. Thebus subsystem may include support structure 520, at least one boosterrocket 522, one or more fuel tanks 526, and one or more pressurizationtanks 528. Support structure 520 is fixed to outer load path supportstructure 502, and physically supports booster rocket 522, fuel tanks526, and pressurization tanks 528. Booster rocket 522 may be operated tochange the altitude of satellite 500 significantly. For example, boosterrocket 522 may be operated to transition satellite 500 from a lowaltitude orbit into an operational orbit (e.g., a Molniya orbit), aswill be described in more detail later. Booster rocket 522 may include,for example, a liquid propellant engine (LPE), such as a bi-propellantliquid rocket adapted to produce high specific impulses. Fuel tanks 526are adapted to contain liquid propellant for use by booster rocket 522and thrusters 524. Pressurization tanks 528 are adapted to maintainadequate pressure within fuel tanks 526.

In an embodiment, satellite 500 also includes a set of multiplethrusters 524 positioned at various locations around satellite 500.Thrusters 524 are adapted to provide moments and velocity for attitudeadjustments, spin up and spin down, and to compensate for orbitalperturbations. Thrusters 524 may be positioned at various positions onsatellite 500 in order to provide these adjustments in a fuel-efficientmanner, as will be described in detail later in conjunction with FIGS.15-18. Although only two thrusters 524 are illustrated in FIG. 5,satellite 500 may include many thrusters, in other embodiments.

Satellite 500 may include a “spinning” bus, in an embodiment. In such anembodiment, the bus subsystem and outer load path support structure 502spin around the primary axis 504, and the communications payloadsubsystem remains substantially stationary, with respect to the primaryaxis 504. In addition, adjustments are made to point the reflector 542of the communications payload subsystem in the correct direction duringoperations. In another embodiment, a satellite may not include aspinning bus (e.g., a body or three axis stabilized satellite).

In an embodiment, the bus subsystem is coupled to the communicationspayload subsystem through a despin control mechanism 530, such as aBearing and Power Transfer Assembly (BAPTA), which is adapted to allowthe bus subsystem to spin around primary axis 504 at a rate wherein thecommunications payload subsystem appears substantially stationary, withrespect to earth. In addition, despin control mechanism 530 is adaptedto provide a conduit for power and signal transmission between the twosubsystems. In other embodiments, a spinning bus may not be implemented,and the bus subsystem and communications payload subsystem may be fixed,relative to each other.

The communications payload subsystem is located substantially withininterior cavity 508, and is exposed at earth opening 510. Thecommunications payload subsystem may include one or more uplink antennas540, a reflector 542, a feed horn support structure 544, a feed horn546, and communications electronics 548, among other things. Inaddition, communications payload subsystem includes a platform 550,which physically supports reflector 542, feed horn support structure544, feed horn 546, and communications electronics 548. In anembodiment, platform 550 includes an inner region 552 and an outerregion 554. Inner region 552 includes a substantially flat and circularstructure, and outer region 554 includes a substantially conicalstructure, which is bisected by inner region 552, in an embodiment.

Uplink antennas 540 are coupled to the periphery of reflector 542, in anembodiment, although they may be coupled to other parts of satellite500, in other embodiments. Uplink antennas 540 receive uplink signals(e.g., signals 110, FIG. 1). In an embodiment, an uplink antenna 540 mayinclude a planar array antenna, although other types of uplink antennasalternatively may be used. Each uplink antenna 540 may receive an uplinksignal of a particular polarization, and may provide the received uplinksignal to communications electronics 548. For example, an uplink signalmay include a dual-polarized signal, and accordingly, two uplinkantennas 540 may be used, where a first uplink antenna 540 receives asignal having a first polarization, and a second uplink antenna 540receives the component of the uplink signal having a secondpolarization. In alternate embodiments, a single antenna may receive adual-polarized signal, or the uplink signal may be singly-polarized, andsatellite 500 may include as few as one uplink antenna 540, or theuplink antenna may be diplexed with reflector 542 such that only asingle antenna is used for both receiving and transmitting.

In an embodiment, communications electronics 548 receive the uplinksignals from uplink antennas 540, and produce downlink signals.Communications electronics 548 may include, for example, a plurality oflow noise amplifiers, downconverters, input filters, traveling wave tubeamplifiers (TWTAs), and output multiplexers, in an embodiment.Communications electronics 548 are mounted on platform 550, in anembodiment. For example, power amplifiers and TWTAs may be mounted onouter region 554, and multiplexers may be mounted on inner region 552,in an embodiment. The various communications electronics 548 elementsmay be mounted in other locations, in other embodiments.

In an embodiment, a downlink signal may include multiple channels oftelevision content, which may be multiplexed, and/or which may be spacedapart over the available spectrum. For example, a downlink signal mayinclude 24 channels, which may be spaced apart by about 40 MHz whentransmitted within the C-band. The active bandwidth per channel may beless than the spacing to mitigate potential effects of interference. Forexample, the active bandwidth per channel may be about 36 MHz. In anembodiment, a modulation system may provide one bit per second perHertz, so that each channel may carry 36 Mbps. In alternate embodiments,a downlink signal may include more or fewer channels, have larger orsmaller channel spacings and/or active bandwidths per channel, and/orcompression algorithms may enable multiple channels to be carried ineach band (e.g., 18 television channels per band) as well as variousalternate modulation and coding schemes. In still other embodiments, oneor more channels may be used to carry other types of information.

The TWTAs are adapted to amplify the downlink signals prior totransmission. In an embodiment, the number of TWTAs is directly relatedto the number of downlink channels provided by satellite 500. Forexample, in an embodiment in which 24 channels are provided on thedownlink, communications electronics 548 may include 24 active TWTAs,each of which produces an output signal within a distinct sub-band. Inaddition, communications electronics 548 may include one or more spareTWTAs, which may be activated during operations, for example, when anactive TWTA malfunctions. The multiplexers combine the signals producedby the TWTAs, and provide the multiplexed signals to one or more antennaports (not illustrated). Each multiplexed signal may include multiplechannels (e.g., 12 channels each).

Feed horn 546 receives the multiplexed signals via the antenna ports,and produces one or more polarized, downlink signals. For example, in apreferred embodiment, a first multiplexed signal may be polarized usinga first circular polarization (e.g., a right hand circularpolarization), and a second multiplexed signal may be polarized using asecond polarization (e.g., a second polarization that is substantiallyorthogonal to the first polarization, such as a left hand circularpolarization). In an alternate embodiments, only a single polarizationmay be implemented or different orthogonal polarizations may beemployed. Feed horn 546 projects the downlink signals toward reflector542, which reflects the downlink signals in a beam direction indicatedgenerally by beam direction vector 560. Reflector 542 may include aparabolic reflector, for example, although other types of reflectorsalternatively may be used. In an embodiment, reflector 542 includes areflector having a shape that produces substantially circular coveragearea contours (e.g., contours 301-306, FIG. 3). In other embodiments,reflector 542 includes a shaped (e.g., distorted) parabolic reflector,where the shape of the reflector produces irregularly shaped coveragearea contours. Reflector 542, feed horn support structure 544, and feedhorn 546 are substantially fixed in orientation, with respect to eachother, and together form a antenna assembly.

The direction beam vector 560 is affected by the orientation of theantenna assembly (i.e., reflector 542, feed horn 546, and feed hornsupport structure 544) with respect to the primary axis 504. Thesignificance of these various axes and the orientation of the antennaassembly are described in more detail in conjunction with FIG. 6.

FIG. 6 illustrates a perspective, cut-away view of a satellite 600, inaccordance with an example embodiment. As mentioned previously, outerload path support structure 602 defines a primary axis 604 of satellite600, and reflector 606 reflects downlink signals in a directionindicated generally by beam direction vector 608. The general directionof beam direction vector 608 is affected by rotation of the antennaassembly about at least two control axes, which include the primary axis604 and an elevation adjustment axis 610, which may be substantiallyorthogonal to the primary axis 604, in an embodiment. For a spinning busembodiment, rotation about primary axis 604 (or “spin axis”) isdesignated as an azimuth adjustment and is effected by a despin controlmechanism (e.g., despin control mechanism 530, FIG. 5), and rotationabout elevation adjustment axis 610 is designated as elevationadjustment and is effected by an elevation adjustment mechanism 620. Inan embodiment, the despin control mechanism (e.g., despin controlmechanism 530, FIG. 5) is adapted to provide a 360 degree rangeadjustment around the primary axis 604, and the elevation adjustmentmechanism 620 is adapted to provide about a 90 degree range adjustmentaround the elevation adjustment axis 610. These two mechanisms areadjusted in accordance with the pointing requirements of a particularorbit plane, such that an antenna pattern which is a figure of rotationmay be maintained over a coverage area. In other words, the at least twocontrol axes may be used to orient the antenna assembly in a directionto support downlink signal transmission into a desired coverage.

FIGS. 7-9 illustrate perspective top views 702, 802, 902 and cut-awayside views 704, 804, 904 of a satellite 700 with a antenna assembly infirst, second, and third positions in elevation with respect to platform710, in accordance with an example embodiment. Referring first to FIG.7, the antenna assembly is positioned, with respect to platform 710, sothat reflector 712 reflects a downlink signal in a direction generallyalong a first beam direction vector 706. Referring now to FIG. 8,rotation of the antenna assembly via elevation adjustment mechanism 810,results in the antenna assembly being positioned, with respect toplatform 710, so that reflector 712 reflects the downlink signal in adirection generally along a first beam direction vector 806. Referringnow to FIG. 9, further rotation of the antenna assembly via elevationadjustment mechanism 810, results in the antenna assembly beingpositioned, with respect to platform 710, so that reflector 712 reflectsthe downlink signal in a direction generally along a first beamdirection vector 906. As FIGS. 7-9 illustrate, the direction of beamdirection vector 706, 806, 906 can be changed significantly, thusproviding the capability to compensate for the pointing requirementsthroughout the operational portion of each orbit, during the operationallife of the satellite 700.

In order to more fully depict the structure and arrangement of asatellite, in accordance with example embodiments, FIGS. 10 and 11 areprovided, which illustrate a perspective view and a top view of asatellite, respectively. Referring first to FIG. 10, satellite 1000 isshown to include outer load path support structure 1002, whichsubstantially contains the communications payload subsystem and the bussubsystem. As discussed previously, outer load path support structure1002 may be substantially cylindrical, in an embodiment, with asubstantially circular cross-section. In other embodiments, outer loadpath support structure 1002 may have a different cross-sectional shape(e.g., square, triangle, pentagon, hexagon, octagon, etc.). Outer loadpath support structure 1002 may be generally referred to as an elongatedsupport structure adapted to contain the communications payloadsubsystem and the bus subsystem. In an embodiment, as will be discussedin conjunction with FIG. 5, outer load path support structure 1002 isadapted to support stacking of one or more other satellites above and/orbelow satellite 1000 within the payload fairing (e.g., payload fairing1904, FIG. 13) of a launch vehicle. In an embodiment, outer load pathsupport structure 1002 is further adapted to support at least one solarpanel 1004 on its exterior surface 1006. Outer load path supportstructure 1002 includes an earth-oriented opening 1008, in anembodiment, which exposes reflector 1010.

Referring to FIG. 11, a top view of a satellite 1100 is illustrated. Asdiscussed previously, outer load path support structure 1102substantially contains the communications payload subsystem and the bussubsystem, and therefore defines an outer surface of satellite 1100. Bussubsystem is fixed to outer load path support structure 1102 with bussubsystem support structure 1104. During operations, when viewed fromthe earth, communications payload subsystem (e.g., reflector 1106, feedhorn 1108, platform 1110, and communications electronics 1112) does notappear to rotate, while outer load path support structure 1102 and bussubsystem (e.g., bus subsystem support structure 1104, fuel tanks 1114,and fuel tanks 1116) do appear to rotate, in an embodiment. As discussedpreviously, the relative rotation between the subsystems is achievedusing a despin control mechanism (e.g., despin control mechanism 530,FIG. 5). In another embodiment, the subsystems may not rotate, relativeto each other.

As discussed previously, a system includes multiple satellites (e.g.,six satellites, or some other number), which may travel within NGSO(e.g., Molniya or other orbits) and/or GSO orbits, and which may have aphasing maintained between each other in a manner that the satellitefleet may provide continuous coverage within one or more coverage areas.In order to achieve such a satellite fleet, each satellite is placed inan appropriate orbital plane and nodal position with respect to theother satellites of the fleet. In an embodiment, implementing thesatellite fleet includes launching at least one set of satellites withina single launch vehicle, and then moving each satellite of the set intoan appropriate orbit position, with respect to the other satellites ofthe fleet.

FIG. 12 illustrates a flowchart of a method for implementing a satellitefleet, in accordance with an example embodiment. As will be described indetail below, embodiments include launching multiple satellites (e.g.,three satellites or some other number) within a single launch vehicleinto an initial orbit, and subsequently transitioning each satelliteinto an operational orbit while achieving a desired nodal separationbetween the satellites. An “initial orbit” may be defined as an orbitinto which the satellites initially are released from a launch vehicle.The initial orbit may be defined by an initial semimajor axis, orbiteccentricity (which in turn defines the altitudes at apogee andperigee), angle of inclination, node crossing (generally termed theright ascension of the ascending node, or RAAN), argument of perigee,and time of perigee (or true anomaly). An “operational orbit” may bedefined as a “final” orbit within which system operations are performed(e.g., a Molniya orbit or other orbit, as discussed previously). Theoperational orbit may be defined by an operational semimajor axis, orbiteccentricity, angle of inclination, node crossing, argument of perigee,and time of perigee.

During the process of transitioning between the initial orbit and theoperational orbit, each satellite may be transitioned to one or moreintermediate orbits, as will be described in detail below. As will alsobe described in detail below, the system may permit each satellite todrift, for various periods of time, at the initial orbit and/or one ormore intermediate orbits, in order to achieve a desired nodal separationbetween the satellite and other satellites of the fleet. Accordingly,embodiments of the inventive subject matter take advantage of theearth's gravitational field, specifically the difference in nodalregression rates between higher and lower orbits, to achieve the nodalseparation between the satellites of a multiple-satellite launch. Thisis in contrast to traditional systems, which achieve nodal separation byfiring booster rockets and consuming limited on-board fuel resources orby using dedicated launch vehicles (i.e., one launch vehicle persatellite). As the below description will convey, embodiments may havethe advantages of economically implementing multiple satellites withintheir operational orbit positions while efficiently conserving limitedfuel resources on-board each satellite.

The method begins, in block 1202, by loading multiple satellites withinthe payload fairing of a launch vehicle, in preparation for launch. Inan embodiment, the launch vehicle includes a launch vehicle (e.g., aLand Launch Zenit-2SLB launch vehicle) capable of injecting the multiplesatellites into the initial orbit. In an embodiment, three satellitesare loaded in a stacked configuration within a payload fairing. Inalternate embodiments, two, four, or another number of satellites may beloaded in a stacked configuration within a payload fairing. The numberof satellites loaded within a single payload fairing may depend, atleast in part, on the size and weight of each satellite, the physicalcapacity of the payload fairing, and the lift capacity of the launchvehicle.

FIGS. 13 and 14 illustrate a cross-sectional, side view and aperspective view, respectively, of a configuration of multiplesatellites 1300, 1301, 1302 stacked within the payload fairing 1304 of alaunch vehicle, in accordance with an example embodiment. In anembodiment, payload fairing 1304 may have a diameter in a range of about2 m to about 5 m, with a diameter of about 3.5 m in a particularembodiment. Satellites 1300-1302 are coupled together at the top and/orbottom edges of outer load support structures 1306, 1307, 1308. Whencoupled together, satellites 1300-1302 may be referred to herein as a“satellite stack” or a “group of satellites.” In an embodiment,satellites 1300-1302 are coupled together with a plurality of explosivebolts and compressed, axial spring mechanisms (not illustrated), whichmay later be activated in an initial orbit to separate satellites1300-1302 from each other. In addition, a spin up platform 1310 isattached to the bottom of lower satellite 1300, in an embodiment. Aswill be described in more detail later, “spin up” rockets (notillustrated) on spin up platform 1310 may be activated once thesatellite stack has separated from payload fairing 1304, in order tospin-stabilize the satellite stack, in an embodiment. In otherembodiments, the spin up rockets may be located elsewhere (e.g. towardsthe middle of the satellite stack), or other thrusters (e.g., radialthrusters 1660 and/or 1664, FIG. 16) may be used to achieve the spin up,or the launch vehicle, if capable, may provide the spin up prior torelease of the satellite stack.

Each outer load path support structure 1306-1308 is adapted to couple,at its top and/or bottom, to one or more other satellites. In addition,each outer load path support structure 1306-1308 is adapted to bear theload from any satellite(s) that may be stacked above it. In anembodiment, substantially all of the weight of any higher satellites isborn by outer load support structures 1306, 1307. Accordingly, the outerload support structure 1308 of the lower satellite 1302 is adapted tobear at least the weight from the center satellite 1301 and the uppersatellite 1300, and the outer load support structure 1307 of the centersatellite 1301 is adapted to bear at least the weight from the uppersatellite 1300, in an embodiment.

Traditionally, satellites are either launched one at a time in adedicated launch vehicle, or multiple satellites are launched togetherusing load path structures substantially different from those describedin conjunction with embodiments of the inventive subject matter.Launching satellites separately is extremely expensive, particularlywhen a satellite system includes a significant number of satellites.However, when multiple satellites are launched together, some type ofload path structure is employed so that the satellites are not damagedduring launch. Traditional methods for multiple satellite launch includesome form of dedicated launch vehicle structure to support multiplesatellites within the payload fairing. These support structures includevarious forms of cradles, dispensers or launch vehicle adapters. Toavoid using a dedicated launch vehicle structure, some traditionalsatellites adapted for multiple satellite launches have employed innerload path support structures such as an inner thrust cylinder, ratherthan the outer load path support structures (e.g., structures 1306-1308)of embodiments of the inventive subject matter. In particular,traditional satellites are designed so that, when stacked within apayload fairing, the load of any higher satellite(s) is born by an innerload path structure that runs through the interior of the satellite.Traditional satellites that employ an inner load path stackedconfiguration typically are held together and then separated from eachother using a mechanical connection and an internal thrust cylinder.

The use of outer load support structures 1306-1308, in accordance withan embodiment, may provide one or more advantages over traditionalsatellite configurations. In particular, the relatively large diameterof the outer load path support structure (e.g., structures 1306-1308)may provide substantially greater lateral stiffness, particularly withina high-vibration launch environment, than is provided by an inner loadpath support structure. In addition, because the outer load path supportstructures (e.g., structures 1306-1308) may have diameters that areslightly less than the diameter of the payload fairing (e.g., payloadfairing 1304), placement or mounting of communications payload subsystemor bus subsystem elements is not affected by interference from the loadpath support structure, as it may be in satellites that include innerload path support structures.

Referring again to FIG. 12, once the satellites have been loaded intothe payload fairing, the launch vehicle is prepared for launch at adesired launch site, in block 1204. In an embodiment, for example, alaunch site may include the Cosmodrome, in Baikonur, Russia, althoughother launch sites may be selected in other embodiments.

Once the launch vehicle is ready and cleared to launch, the launchphases may be performed, in block 1206. The launch phases serve to boostthe launch vehicle from the launch platform, and to deliver thesatellite stack (i.e., satellites 1300-1302, FIG. 13) to an initialorbit. In an embodiment, the initial orbit may be substantiallycircular, and the initial altitude is an altitude in a range from 200 kmto 400 km, with an altitude of 300 km being preferred.

Before payload separation (i.e., release of the satellite stack from thepayload fairing), the launch vehicle may impart a slow rotation to thesatellite stack, which may result in settling the fuel in thesatellites' fuel tanks (e.g., fuel tanks 1542, 1543, FIG. 15). Once thesatellite stack has been released from the payload fairing, rockets onthe spin up platform (e.g., spin up platform 1310, FIG. 13) may beactivated to spin up the satellite stack, in block 1208, in anembodiment. In an embodiment, activation of the spin up platform is usedto impart an axial rotation to the satellite stack around thesatellites' primary axes (e.g., primary axis 504, FIG. 5) at a spinvelocity in a range of about 20 to 40 revolutions/minute (rev/min), witha spin velocity of about 30 rev/min being preferred. Once the desiredspin velocity is achieved, the spin up platform (e.g., spin up platform1310, FIG. 13) may be separated from the satellite stack. In anotherembodiment, satellite thrusters (e.g., thrusters 1504-1515, FIG. 15) maybe used to spin up the satellites and/or the satellite stack, in whichcase the spin up platform may be excluded. In still another embodiment,the launch vehicle may spin up the satellite stack prior to release fromthe payload fairing. At that point, the satellite stack may be releasedfrom the payload fairing, or each satellite may be sequentially releasedfrom the payload fairing one at a time. In the latter case, block 1210(described below) may be excluded. Because the satellites are induced tospin around the primary axis, the primary axis may also be referred toas the “spin axis.”

In block 1210, the satellites (e.g., satellites 1300-1302) may beseparated from each other. In an embodiment, the satellites areseparated by detonating explosive bolts that were used to coupletogether the outer load path support structures (e.g., outer load pathsupport structures 1306-1308, FIG. 13) of the satellites. Compressedaxial, separation springs positioned between the outer load path supportstructures may then be freed to impart a relatively small axial velocitythe satellites away from each other. As mentioned previously, thesatellites already may be spinning around their primary axes at anominal speed as a result of impulses generated by the spin up platform(e.g., spin up platform 1310, FIG. 13). In another embodiment, thesatellites may spin themselves up to a nominal spin speed (e.g., about30 rev/min), using appropriate thrusters (e.g., radial thrusters 1660,1664, FIG. 16). In such an embodiment, a spin up platform may beexcluded from the system.

Once the satellites are separated, each satellite may be transitionedfrom the initial orbit to a “parking orbit,” in block 1212. In anembodiment, the parking orbit may be selected to provide for relativelylong communication passes between a satellite and a ground controlstation, in addition to providing substantially faster nodal regressionthan a final operational orbit. During these communication passes, thesatellite may transmit telemetry information to the ground controlstation, and the ground control station may transmit command messages tothe satellite. In an embodiment, the parking orbit may have an apogeealtitude in a range of about 1200-1800 km, with about 1500 km being apreferred apogee altitude. Higher or lower apogee altitudes may be used,in other embodiments. In still other embodiments, where communicationsat a lower orbit such as the initial orbit are not an issue, block 1212may be excluded.

The transition from the initial orbit to the parking orbit may require avelocity increment in a range of about 1500 m/s to 4000 m/sec, in anembodiment, with a velocity of about 2500 m/sec being reasonable. In anembodiment, some or all orbit changes are made using booster rocketburns when a satellite is in the vicinity of perigee. Accordingly, priorto transitioning the satellites from the initial orbit, the satellitesmay be permitted to travel in their orbits until they are at or nearperigee. In an embodiment, each satellite is transitioned to the parkingorbit within a small number of orbital periods after being released atthe initial orbit. In another embodiment, the launch vehicle may placethe satellites directly into the parking orbit, and block 1212 may beexcluded.

After transition into the parking orbit in block 1212 (or after releaseinto the initial orbit, if block 1212 is not performed), the satellitesare separately transitioned into their operational orbits, in anembodiment. To initiate the transition of the satellites to theiroperational orbits, a first satellite is selected for transition, inblock 1214. As will be described below, the selected satellite may betransitioned from the parking orbit (or from the initial orbit) to theoperational orbit by performing a series of one or more orbit changes.Each of these orbit changes may be made, for example, by performing abooster rocket burn (e.g., booster rocket 522, FIG. 5). In anembodiment, prior to beginning the orbit transition, the system maymaintain a satellite in the parking orbit (or initial orbit if block1212 is omitted) for a drift time period, in block 1216, in order toallow the selected satellite to drift toward the desired nodal positionand/or to establish a desired nodal separation between the selectedsatellite and previously transitioned satellites, if any. When a firstsatellite of the fleet has been selected and is being transitioned tothe operational altitude, the system may bypass the process of block1216, as the first satellite's position may be used as the baseline bywhich later-transitioned satellites' nodal positions are determined. Asis described below, embodiments of the inventive subject matter takeadvantage of the differences in nodal regression rates for differentorbits in order to achieve desired nodal separations between satellitesof a fleet. Accordingly, embodiments may use significantly less on-boardfuel resources in order to achieve desired nodal separations thantraditional systems, which rely on booster rocket burns to achieve nodalseparations, or which avoid the use of separate launch vehicles toachieve nodal separations.

The RAAN for a satellite drifts at a rate that varies based on thesatellite's orbit. In particular, the RAAN drifts relatively quickly atthe initial orbit (e.g., approximately 4°/day at 300 km circular), andrelatively slowly at the operational orbit (e.g., approximately0.14°/day at 39,560 km). The rate at which a RAAN drifts may be referredto as a nodal regression rate. As mentioned above, a selected satellitemay be maintained at one or more orbits that are lower than theoperational orbit (e.g., the initial orbit, the parking orbit, and/orone or more other intermediate orbits) for one or more drift timeperiods, in order to allow the selected satellite to drift and establisha desired nodal separation between the selected satellite and one ormore other satellites of the fleet. For example, the system may wait acumulative drift time period in a range of about 12 days to about 20days to achieve a 60° nodal separation between neighboring satelliteswithin a fleet, in an embodiment. Alternatively, the system may waitlonger or shorter time periods to achieve larger or smaller nodalseparations, respectively. In some cases, a portion of the desired nodalseparation may occur while the satellite is being transitioned betweenorbits.

During the transition to the operational orbit, the selected satellite'sattitude and orbit repeatedly may be adjusted. For example, in order toachieve fuel-efficient booster rocket burns, the selected satellite'sattitude may be adjusted prior to each burn, and each burn maytransition the satellite only partially toward the operational orbit.The iterative process of transitioning each satellite to its operationalorbit is conveyed in blocks 1218-1222.

In block 1218, the selected satellite may perform one or more attitudeadjustments. Because of the orientation of the booster rocket (e.g.,booster rocket 522, FIG. 5) with respect to the satellite's primary axis(e.g., primary axis 504, FIG. 5), the thrust direction is along thesatellite's primary axis, in an embodiment. Attitude adjustmentsdesirably are made so that, prior to an orbit change (e.g., prior to abooster rocket burn), the satellite's primary axis is pointed in adirection that may result in the desired orbit change using a minimalamount of fuel. Using telemetry information received from the satellite(and/or other information) during a communication pass, the groundcontrol station may determine the satellite's current attitude (e.g.,current roll, pitch, and yaw angles). The ground control station maythen compare the current attitude to a desired attitude (e.g., a desiredattitude that should be attained prior to a next rocket booster burn),and determine parameters for thruster activation (or other means ofchanging the satellite's attitude) that will enable the satellite toachieve the desired attitude. For example, thruster activationparameters may indicate the identities of thrusters that should beactivated, start and stop times (or a duration) for each thrusteractivation, thrust levels, and thruster activation sequence information.The ground control station may then generate one or more attitudeadjustment command messages that indicate the thruster activationparameters, and send the command messages to the satellite during thesame communication pass or during a later communication pass. Thesatellite may then perform the attitude adjustments specified in theground control station command messages. In an embodiment, a satellitemay perform one or more attitude adjustment iterations prior toperforming an orbit change.

In block 1220, the satellite may perform one or more orbit changes. Inan embodiment, the ground control station also may use informationreceived from the satellite (and/or other information) to determine thesatellite's current orbit. The ground control station may then comparethe current orbit to a desired orbit (e.g., a next intermediate orbit orthe operational orbit), and determine parameters for a booster rocketburn that will enable the satellite to achieve the desired orbit. Forexample, booster rocket burn parameters may include one or more startand stop times (or durations) for one or more burns. The ground controlstation may then generate one or more orbit adjustment command messagesthat indicate the booster rocket burn parameters, and send the commandmessages to the satellite during the same communication pass or during alater communication pass. The satellite may then perform the orbitadjustments specified in the ground control station command messages.The satellite may make an orbit adjustment by imparting an adequatevelocity increment using a booster rocket (e.g., rocket 522, FIG. 5). Inan embodiment, the ground control station may send separate commandmessages indicating thruster activation parameters and booster rocketburn parameters. In another embodiment, the thruster activationparameters and the booster rocket burn parameters may be sent to asatellite in the same message.

A determination may be made, in block 1222, whether the satellite hasachieved its desired operational orbit. Depending on various factors, itmay take about 7 to 10 days to transition a satellite into its desiredoperational orbit. Alternatively, the transition may take a longer orshorter period of time. When the satellite has not achieved it desiredoperational orbit, then the method may iterate as shown.

When the satellite has achieved the desired operational orbit, thesatellite is prepared for system operations, in block 1224. In anembodiment, this includes adjusting the satellite's attitude to placethe satellite in an operational attitude. In an embodiment, as discussedpreviously, the operational attitude is an ecliptic normal attitude(i.e., the satellite's primary axis (e.g., primary axis 504, FIG. 5) isperpendicular to the ecliptic). Adjustment to the operational attitudemay be achieved through thruster activations, as discussed previously.Preparing the satellite for system operations also may include“de-spinning” the communications payload subsystem, in an embodiment.This may be achieved by activating the despin control mechanism (e.g.,despin control mechanism 530, FIG. 5) and damping the spin on thecommunications payload subsystem until it appears relatively stationary,when viewed from the earth.

Preparation for system operations also may include orienting thesatellite's antenna assembly in the proper direction to support uplinkreception and downlink signal transmission into a desired coverage area(e.g., coverage area 306, FIG. 3). As described previously, this mayinclude orienting the antenna assembly about an azimuth axis (e.g., axis604, FIG. 6, as controlled by the despin control mechanism 530, FIG. 5)and an elevation axis (e.g., axis 610, FIG. 6, as controlled by theelevation adjustment mechanism 620). These adjustments, together with afigure of revolution antenna pattern, allow correct antenna pointingover a coverage area and compensate for differences in orbit planes(RAAN) and nodal regression, as discussed previously. These rotationaladjustments may be performed continuously or in periodic, discrete stepsduring the operational life of the satellite.

In block 1226, a determination may be made whether all satellitespreviously in the satellite stack have been transitioned into thedesired operational altitude. When all satellites have not beentransitioned, then another satellite is selected, in block 1214, and themethod iterates as shown for the newly selected satellite.

When all satellites have been transitioned into their desiredoperational orbits, the method ends for this set of satellites. When asatellite fleet includes more than the number of satellites within thefirst launch (e.g., more than three satellites), as is the case forsystem embodiments previously described herein, the entire process maybe repeated for one or more additional groups of satellites. In a systemembodiment that includes six satellites within its fleet, the processmay be repeated once. In other words, a second launch vehicle may belaunched, which includes a second group of satellites within its payloadfairing. The second group may be released in an initial orbit, and eachsatellite of the second group may be sequentially transitioned to theiroperational orbits, in a manner such that desired nodal separations areestablished between satellites of the first group and satellites of thesecond group. In an embodiment, the initial orbit for the second groupof satellites may be about 180° separated from the initial orbit for thefirst group of satellites. In other embodiments, the initial orbit forthe second group of satellites may have a different separation anglefrom the initial orbit for the first group of satellites. Once thesatellites within the fleet have been successfully positioned withintheir operational orbits, regular system operations may be performed.For example, the system may begin providing satellite services (e.g.,DBS television services), in an embodiment.

Example embodiments described above discuss launching three satelliteswithin a single launch vehicle. Those of skill in the art wouldunderstand, based on the description herein, how to alter the abovedescribed embodiments in order to launch more (e.g., four or more) orfewer (e.g., two) satellites within a single launch vehicle, and furtherhow to transition those satellites into their operational orbits whileachieving desired nodal separations. Accordingly, alternate embodimentsin which more or fewer satellites are launched within a launch vehicleare intended to be included within the scope of the inventive subjectmatter.

Once the satellite fleet is implemented, the system may beginoperations. In an embodiment, each satellite may be in an inactive modeprior to system operations startup and whenever a satellite is locatedwithin an inactive orbit segment. In an inactive mode, a satellite mayconserve power by maintaining communications and control systems in alow-power state, and the satellite may refrain from transmittingdownlink signals or receiving uplink signals, with the possibleexception of health, status, and control-types of signals. Whenever asatellite is located within an active orbit segment, a satellite may bein an active mode. In an active mode, a satellite may maintaincommunications and control systems in an operational-power state, andthe satellite may receive uplink signals and transmit downlink signals(e.g., television signals).

During the operational life of the satellite, satellite orbitmaintenance is actively performed, in order to correct for orbitaldeviations. The primary orbital parameters that should be controlledinclude the argument of perigee, the semi-major axis, the eccentricity,the inclination and the node (RAAN). Orbital deviations may be caused byhigher order terms of the earth's gravitational field, gravitationalfields of the sun and the moon, and other factors, such as solarpressure. These perturbing forces may be countered by activating thesatellite thrusters (e.g., thrusters 524. FIG. 5) and/or the satellitebooster rocket (e.g., booster rocket 522, FIG. 5). Both during and afterimplementation of the satellite fleet, the satellite thrusters may beused to perform spin up operations (e.g., process 1208, FIG. 12), spindown operations, and attitude adjustments (e.g., process 1218, FIG. 12),while the booster rocket may be used to perform significant orbitchanges (e.g., process 1220, FIG. 12).

In an embodiment, velocity increments to maintain the satellite's orbitmay be in the orbit plane or nearly perpendicular to the orbit plane.For example, velocity increments in the orbit plane may be applied tocorrect semi-major axis, eccentricity, and argument of perigee orbitalerrors, in an embodiment. Velocity increments perpendicular to the orbitplane may be applied to correct inclination and RAAN. To correct theorbital errors, thrusters are placed on a satellite in an arrangementwhere velocity increments may be provided in three perpendicular axes(or nearly perpendicular axes), in an embodiment. Orbital errors may becorrected by firing one or more thrusters at appropriate times in theorbit. For embodiments that include spinning satellites, the thrusterfirings are made at the appropriate times in the spin cycle. Because thesatellites are operated in ecliptic normal attitudes, in an embodiment,the orbit plane varies with respect to the coordinate system of thesatellite. Accordingly, when the thruster firings do not result in thedesired in-plane or perpendicular velocity increments, then thethrusters may be fired at appropriate times and for appropriatedurations and vector summed to achieve the desired increment. Thisresults in a significant total velocity increment that may be applied inorder to maintain the satellite's orbit. For example, a total velocityincrement in a range of about 1000 m/s to about 2000 m/s may be applied,with a total velocity increment of about 1500 m/s being possible in anembodiment.

Embodiments of the inventive subject matter include methods andapparatus for providing satellite orbit maintenance in a manner that mayconsume on-board fuel in an efficient and conservative manner. In anembodiment, a propulsion subsystem is provided, which includes a boosterrocket and a plurality of thrusters strategically positioned around thesatellite so that on-board fuel is efficiently consumed. Embodiments ofpropulsion subsystems and thruster positions are illustrated anddescribed in conjunction with FIGS. 15-17.

FIG. 15 illustrates a simplified diagram of a propulsion subsystem 1500,in accordance with an example embodiment. Propulsion subsystem 1500includes a booster rocket 1502 (e.g., an LPE), a plurality of axialthrusters 1504, 1505, 1506, 1507, a plurality of radial thrusters 1508,1509, 1510, 1511, and a plurality of canted thrusters 1512, 1513, 1514,1515, in an embodiment. Booster rocket 1502 and thrusters 1504-1515 areillustrated as hypergolic chemical thrusters, but may include, forexample, Hall-effect thrusters, plasma thrusters, electric thrusters,ion engines, and/or other types of thrusters adapted to produce a highspecific impulse. Although four each of axial thrusters 1504-1507,radial thrusters 1508-1511, and canted thrusters 1512-1515 (i.e., twelvetotal thrusters 1504-1515) are illustrated in FIG. 15, propulsionsubsystem 1500 may include more and/or fewer of any particular type ofthrusters. Further, in various embodiments, certain ones of thrusters1504-1515 may be primary thrusters (i.e., thrusters used during normaloperations), and certain ones of thrusters 1504-1515 may be backupthrusters (e.g., in the event of failure of a primary thruster). In somecases, a particular thruster may serve as both a primary and a backupthruster.

Propulsion subsystem 1500 also includes one or more fuel tanks 1520,1522 and one or more oxidizer tanks 1524, 1526, which are coupled tobooster rocket 1502 and thrusters 1504-1515 through supply lines 1528,1530. Propulsion system 1500 may also include a pressurization system1540 to maintain adequate pressure within fuel tanks 1520, 1522 andoxidizer tanks 1524, 1526, and may also include valves, regulators,filters, and other components (not illustrated). Thrusters 1504-1515 maybe placed in various locations on a satellite in order to provide forsatellite attitude, altitude, and primary axis spin adjustments.Thrusters 1504-1515 may be fired in time bursts and/or in pulsedsequences. Embodiments of thruster placement are illustrated anddescribed in conjunction with FIGS. 16 and 17.

FIG. 16 illustrates a simplified bottom view of a satellite 1600,illustrating rocket booster 1602, axial thrusters 1604, 1605, 1606,1607, and radial thrusters 1608, 1609, 1610, 1611, in accordance with anexample embodiment. Rocket booster 1602 may be positioned in the centerof satellite 1600, so that it may impart an impulse coincident with theprimary axis (e.g., primary axis 504, FIG. 5) of satellite 1600 (i.e.,in a direction that is into the page). Accordingly, rocket booster 1602is adapted to move satellite 1600 from one orbit to another (e.g.,raising the orbit apogee).

Axial thrusters 1604-1607 may be positioned at the bottom of satellite1600 between the primary axis and the outer cylinder 1650 of satellite1600. In an embodiment, axial thrusters 1604-1607 are positioned in arange of 50% to 90% of the distance from the primary axis to the outercylinder 1650 of the satellite. In other embodiments, axial thrusters1604-1607 may be positioned closer to or farther from the primary axis.Axial thrusters 1604-1607 are oriented to impart impulses that aresubstantially parallel to the primary axis (i.e., in a direction that isinto the page). When fired simultaneously, axial thrusters 1604-1607also may impart an impulse coincident with the primary axis.Accordingly, axial thrusters 1604-1607 are adapted to provide a velocityincrement in the axial direction to control the orbit. However, whenfired separately and/or sequentially, axial thrusters 1604-1607 maycause satellite 1600 to rotate about axes (not illustrated), which areperpendicular to the primary axis. Accordingly, axial thrusters1604-1607 are adapted to re-orient the satellite's primary axis. In anembodiment, a primary thruster for primary axis control may be any oneor more of axial thrusters 1604-1607. If a primary thruster fails, thenanother one or more of axial thrusters 1604-1607 may be consideredbackup thrusters for primary axis adjustment.

Radial thrusters 1608-1611 may be positioned around the outer cylinder1650 of satellite 1600. Radial thrusters 1608-1611 are oriented toimpart impulses that are substantially perpendicular to the primaryaxis. In an embodiment, radial thrusters 1608-1611 are located inpositions that are coincident with a plane that bisects the satellite'scenter of gravity (e.g., plane 1714, FIG. 17), and which isperpendicular to the primary axis. Radial thrusters 1608-1611 may bepositioned at substantially equal distances from each other around outersurface 1650 (e.g., approximately 90° apart around a substantiallycircular outer surface), in an embodiment. In other embodiments, thedistances between radial thrusters 1608-1611 may be unequal.

In an embodiment, a first set of radial thrusters, which includes atleast one radial thruster (e.g., radial thrusters 1608, 1610), mayimpart impulses, indicted by impulse vectors 1660, which produce aclockwise rotational vector 1662 around the primary axis. A second setof radial thrusters, which includes at least one other radial thruster(e.g., radial thrusters 1609, 1611), may impart impulses, indicated byimpulse vectors 1664, which produce a counter-clockwise rotationalvector 1666 around primary axis. Accordingly, radial thrusters 1608-1611are adapted to increase and decrease a rate of spin around the primaryaxis. In an embodiment, the first set of radial thrusters includesradial thrusters 1608, 1610, and may be the primary set of thrusters forspin up of satellite 1600. If either of radial thrusters 1608, 1610fails, then the other one of radial thrusters 1608, 1610 may beconsidered the backup thruster for spin up of satellite 1600. The secondset of radial thrusters includes radial thrusters 1609, 1611 may be theprimary set of thrusters for spin down of satellite 1600. If either ofradial thrusters 1609, 1611 fails, then the other one of radialthrusters 1609, 1611 may be considered the backup thruster for spin downof satellite 1600. When fired in pairs, such as 1608 and 1611 or 1609and 1610, the radial thrusters impart a radial velocity (i.e., avelocity perpendicular to the spin axis) on the satellite 1600. Thethrusters are pulsed and the direction of the velocity is determined bythe position of the thruster within the spin cycle when the pulseoccurs.

FIG. 17 illustrates a simplified side view of a satellite 1700,illustrating canted thrusters 1702, 1703, 1704, 1705, in accordance withan example embodiment. Canted thrusters 1702-1705 may be positionedaround the outer surface 1710 of satellite 1700. Canted thrusters1702-1705 are oriented to impart impulses that substantially intersectthe satellite's center of gravity, as indicated by vectors 1712, 1713,1714, 1715. Canted thrusters 1702-1705 are adapted to provide a velocityincrement. As with the radial thrusters, the canted thrusters may bepulsed, and the direction of the velocity increment is determined by thethruster selected and the position of the thruster within the spincycle. Multiple canted thrusters (e.g., thrusters 1702-1705), axialthrusters (e.g., thrusters 1604-1607, FIG. 16), and radial thrusters(e.g., thrusters 1608-1611, FIG. 16) may move satellite 1700 in adirection coincident with the sum of the impulse vectors from thethrusters. By appropriate selection of the thrusters fired and theirdurations, the fuel required to provide orbital correction is reducedcompared to conventional orthogonal thruster sets.

In an embodiment, a first set of “upper” canted thrusters, whichincludes at least one canted thruster (e.g., canted thrusters 1702,1703), may be positioned above a plane 1716 that bisects the satellite'scenter of gravity, and which is perpendicular to primary axis 1718. Asecond set of “lower” canted thrusters, which includes at least onecanted thruster (e.g., canted thrusters 1704, 1705), may be positionedbelow plane 1716. A “cant angle” may be defined as an angle between athruster's impulse vector (e.g., vectors 1712-1715) and plane 1716. Inan embodiment, the cant angle for upper canted thrusters 1702, 1703 maybe in a range of about 30° to about 60°, with a cant angle of about 44°being preferred. The cant angle for lower canted thrusters 1704, 1705may be in a range of about 30° to about 60°, with a cant angle of about38.4° being preferred.

A first subset of the available axial thrusters (e.g., thrusters1604-1607, FIG. 16), radial thrusters (e.g., thrusters 1608-1611),and/or canted thrusters 1702-1705 may be considered the primarythrusters for satellite orbit maintenance, in an embodiment. Forexample, axial thrusters 1605, 1607, radial thrusters 1609, 1610, andcanted thrusters 1703, 1704 may be considered the primary thrusters forsatellite orbit maintenance. If any one or more of the primary thrustersfails, then one or more of the other axial, radial, and/or cantedthrusters may be used as backup thrusters.

FIG. 18 illustrates a simplified block diagram of various satellitesubsystems, in accordance with an example embodiment. Although thedescription of FIG. 18 is directed toward a satellite that provides DBSsatellite services, a satellite that implements embodiments of theinventive subject matter may provide other commercial or non-commercialservices, in other embodiments. Satellite 1800 may includepropulsion/orbit maintenance/attitude control subsystems 1802, a powersubsystem 1804, at least one communications subsystem 1806, an uplinkantenna subsystem 1808, a downlink antenna subsystem 1810, in anembodiment. Power subsystem 1804 may provide electrical power topropulsion/orbit maintenance/attitude control subsystems 1802,communications subsystem 1806, uplink antenna subsystem 1808, anddownlink antenna subsystem 1810, in an embodiment. Power subsystem 1804may include one or more solar cell assemblies (including one or moresolar panels attached to an outer structure, such as the outer load pathsupport structure), and one or more batteries, for example.

Uplink antenna subsystem 1808 may receive uplink signals (e.g., uplinktelevision signals) transmitted from an uplink hub (e.g., hub 104, FIG.1). In an embodiment, uplink signals include uplink television signals,which may include multiple channels of content (e.g., audio and video)that are multiplexed or otherwise combined together. Uplink signals mayinclude other types of information, in addition or alternatively. In anembodiment, an uplink receive antenna may include a circularly-polarizedor a linearly-polarized antenna. In either case, dual (orthogonal)polarizations may be used in order to make full use of the allocatedspectrum.

Communications subsystem 1806 may channelize and amplify uplinkcommunication signals, downconvert the channels to the transmitfrequency band, and amplifies the signals for transmission at theappropriate power level to the ground, in an embodiment. Final stageamplifiers may include an active traveling wave tube for each RFchannel. Final stage amplifiers may include solid state poweramplifiers, in other embodiments. In an embodiment, twelve channels foreach polarization may be multiplexed together to provide the inputs tothe downlink antenna.

In addition, communications subsystem 1806 may receive uplink controlsignals regarding operation of the propulsion/orbit maintenance/attitudecontrol subsystems 1802 and the communications subsystem 1806, amongother things, in an embodiment. In response to the uplink controlsignals, communications subsystem 1806 may be used to activate thetransmitters and/or receivers during an active orbit segment as analternative to an autonomous on-board system for achieving thisfunction. In addition, communications subsystem 1806 may sendinformation to propulsion/orbit maintenance/attitude control subsystems1802 regarding firing sequences for propulsion subsystem elements. In anembodiment, communications subsystem 1806 may transmit appropriateinformation (e.g., health/status and telemetry information) to a controlstation, which may be used in managing the operation of the satellite.

Downlink antenna subsystem 1810 transmits communication channels to theground (e.g., 24 communication channels) with a directional pattern thatcontains all of the user equipment systems. In an embodiment, theantenna beam is pointed in the appropriate direction using a despincontrol system and a mechanism that rotates the antenna in elevation.Other embodiments may include an additional mechanism to rotate theantenna about the boresight. In an embodiment, downlink antennasubsystem 1810 includes a downlink antenna adapted to produce an antennapattern that is a figure of revolution.

Propulsion/orbit maintenance/attitude control subsystems 1802 may beadapted to provide velocity increments for orbit changes, maintainsatellite 1800 in the satellite's orbit, and make satellite attitudeadjustments, in an embodiment. Although propulsion, orbit maintenance,and attitude control may be performed by distinct subsystems, in anembodiment, they may share some common system elements (e.g., thrusters1504-1515, FIG. 15). Accordingly, for purposes of explanation, they areillustrated in a single block in FIG. 18. Alternatively, they may bereferred to separately as a propulsion subsystem, an orbit maintenancesubsystem, and an attitude control subsystem.

In order to provide orbit maintenance and/or attitude control,propulsion/orbit maintenance/attitude control subsystems 1802 maycalculate and/or receive attitude and/or orbit adjustment information,and may operate one or more propulsion subsystem elements in order toadjust the satellite's attitude and/or orbit. In a particularembodiment, propulsion/orbit maintenance/attitude control subsystems1802 are adapted to maintain satellite 1800 in a highly elliptical orbit(e.g., a Molniya orbit) with an orbital period of about 12 hours and anangle of inclination of about 63.4 degrees, while countering theperturbing gravitational influences of the earth, sun, and moon. Asdiscussed previously, satellite 1800 may be maintained in other NGSOand/or GSO orbits, in other embodiments. Maintaining the orbit may alsoinclude maintaining equal nodal separations between the satellites andother satellites within a satellite fleet (e.g., nodal separation ofabout 60 degrees for a fleet of six satellites). In a particularembodiment, propulsion/orbit maintenance/attitude control subsystems1802 may be adapted to maintain an orbit phasing of the satellite, withrespect to other satellites within the satellite fleet, so that thesatellite enters an active orbit segment and initiates transmission ofdownlink signals as a second satellite exits an active orbit segment andceases transmission of the downlink signals. In an embodiment, orbitalphasing may be adjusted to permit uninterrupted transmissions while thesatellite is within an active orbit segment. In addition,propulsion/orbit maintenance/attitude control subsystems 1802 may beadapted to maintain the attitude of satellite 1800.

As discussed previously, satellite 1800 may be maintained in an eclipticnormal attitude, in an embodiment. In a particular embodiment,propulsion/orbit maintenance/attitude control subsystems 1802 may beadapted to maintain satellite 1800 in an ecliptic normal attitude sothat the satellite's outer structure (e.g., substantially cylindricalouter load path support structure 1002, FIG. 10) and the solar arraypanels are approximately perpendicular to the sun throughout the year.In an embodiment, the payload (e.g., power subsystem 1804,communications subsystem 1806, uplink antenna subsystem 1808, anddownlink antenna subsystem 1810) are substantially surrounded by theouter structure, which may result in a thermal boundary that isapproximately constant throughout each orbit and throughout theoperational life of satellite 1800. One or more fuel sources may beassociated with propulsion/orbit maintenance/attitude control subsystems1802, which may be drawn upon by the various propulsion elements.

While at least one exemplary embodiment has been presented in theforegoing detailed description, it should be appreciated that a vastnumber of variations exist. It should also be appreciated that theexemplary embodiment or exemplary embodiments are only examples, and arenot intended to limit the scope, applicability, or configuration of thedescribed embodiments in any way. Rather, the foregoing detaileddescription will provide those skilled in the art with a convenient roadmap for implementing the exemplary embodiment or exemplary embodiments.It should be understood that various changes can be made in the functionand arrangement of elements without departing from the scope as setforth in the appended claims and the legal equivalents thereof.

What is claimed is:
 1. A satellite stack including at least onesatellite, the stack comprising: a first satellite; a first outer loadpath support structure positioned on the first satellite, the outer loadpath support structure comprising a single substantially cylindricalouter load path support structure characterized by an outer radius andhaving an interior cavity adapted to contain at least one satellitesubsystem, the outer load path structure positioned at the outer radiusof the first satellite, a second satellite; a second outer load pathsupport structure positioned on the second satellite; and wherein theouter load path support structure of the first satellite is coupled tothe second outer load path support structure of the second satellite andto substantially bear a load of the second satellite, when the secondsatellite is stacked above the first satellite, a contact between theouter load path support structure of the first satellite and the outerload path support structure of the second satellite being substantiallycontinuous along the outer radius of the first outer load path supportstructure and the second outer load path support structure.
 2. Thesatellite stack of claim 1, further comprising: one or more solar cellassemblies attached to the outer load path support structure of thefirst satellite.
 3. The satellite stack of claim 1, the first satellitefurther comprising: a payload subsystem; and a bus subsystem coupled tothe payload subsystem and to the outer load path support structure,wherein the bus subsystem includes at least one booster rocket and oneor more fuel tanks.
 4. The satellite stack of claim 3, wherein the outerload path support structure of the first satellite is rigidly coupled tothe bus subsystem, and the first satellite further comprises: a despincontrol mechanism adapted to couple the bus subsystem to the payloadsubsystem, wherein the despin control mechanism is further adapted toallow the bus subsystem to spin around a primary axis of the firstsatellite at a rate wherein the payload subsystem appears substantiallystationary, with respect to earth.
 5. The satellite stack of claim 1,the first satellite further comprising: a plurality of axial thrusters,radial thrusters, and canted thrusters adapted to perform satelliteorbit maintenance to correct for orbital deviations.
 6. The satellitestack of claim 1, wherein the second satellite comprises a boosterrocket extending beyond the outer load path support structure of thesecond satellite, the second satellite stacked on top of the firstsatellite, and wherein the inner cavity of the first satellite isadapted to receive the booster rocket of the second satellite while theouter load path support structure of the first satellite substantiallycontacts the outer load path support structure of the second satellite.7. A stack of satellites comprising: a first satellite comprising: apayload subsystem; a bus subsystem coupled to the payload subsystem; afirst outer structure that substantially surrounds the payload subsystemand the bus subsystem, wherein the outer structure comprises a firstouter load path support structure that includes a single substantiallycylindrical structure positioned at a single outer radius of the firstsatellite; one or more solar array panels attached to the outerstructure; and an attitude control subsystem adapted to maintain thefirst satellite in an ecliptic normal attitude during an operationallife of the first satellite so that the one or more solar array panelsare approximately perpendicular to the sun throughout a year, resultingin a thermal boundary that is approximately constant throughout theoperational life of the first satellite; a second satellite comprising asecond outer load path support structure comprising a singlesubstantially cylindrical structure positioned at a single outer radiusof the second satellite; wherein the first outer load path supportstructure is coupled to the second outer load path support structure ofthe second satellite and substantially bears a load of the secondsatellite, the second satellite being stacked above the first satellite,a contact between the first outer load path support structure of thefirst satellite and the second outer load path support structure of thesecond satellite being substantially continuous along the outer radiusof the first outer load path support structure and the outer radius ofthe second outer load path support structure; and wherein the secondsatellite comprises a booster rocket extending below a bottom edge ofthe second outer load path support structure of the second satellite,the second satellite stacked on top of the first satellite, and whereinthe inner cavity of the first satellite is adapted to receive thebooster rocket of the second satellite below an upper edge of the firstouter load path support structure while the first outer load pathsupport structure of the first satellite substantially contacts thesecond outer load path support structure of the second satellite.
 8. Thestack of satellites of claim 7, wherein the attitude control subsystemincludes a plurality of axial thrusters, radial thrusters, and cantedthrusters.
 9. The stack of satellites of claim 7, further comprising:one or more mechanisms adapted to point an antenna beam produced by thefirst satellite in a desired direction.
 10. The stack of satellites ofclaim 7, further comprising: a despin control mechanism adapted tocouple the bus subsystem to the payload subsystem, wherein the despincontrol mechanism is further adapted to allow the bus subsystem to spinaround a primary axis of the first satellite at a rate wherein thepayload subsystem appears substantially stationary, with respect toearth.
 11. The stack of satellites of claim 7, further comprising: anorbit maintenance subsystem adapted to maintain the first satellite in aMolniya orbit with an orbital period of about 12 hours, an angle ofinclination of about 63.4 degrees, and a desired nodal separation of thefirst satellite, with respect to other satellites within a satellitefleet, of 360 degrees divided by a total number of satellites in thesatellite fleet.
 12. A stack of satellites, with one satellite having aprimary axis and an outer cylinder, the stack of satellites comprising:a first satellite comprising: a payload subsystem; a bus subsystemcoupled to the payload subsystem, wherein the bus subsystem includes atleast one booster rocket and one or more fuel tanks; at least onebooster rocket adapted to impart impulses substantially parallel to theprimary axis; an orbit maintenance subsystem adapted to correct fororbital deviations during an operational life of the first satellite,wherein the orbit maintenance subsystem includes a plurality of axialthrusters, radial thrusters, and canted thrusters; a first outer loadpath support structure that substantially surrounds the payloadsubsystem, the bus system, and the orbit maintenance system, the firstouter load path support structure comprising a substantially cylindricalstructure positioned at the outer cylinder of the first satellite; asecond satellite comprising a second outer load path support structurecomprising a single substantially cylindrical structure positioned at asingle outer radius of the second satellite; wherein the first outerload path support structure is couple to the second outer load pathsupport structure of the second satellite and substantially bears a loadof the second satellite, the second satellite being stacked above thefirst satellite, a contact between the first outer load path supportstructure of the first satellite and the second outer load path supportstructure of the second satellite being substantially continuous alongthe outer radius of the first outer load path support structure and theouter radius of the second outer load path support structure; andwherein the second satellite comprises a booster rocket extending belowa bottom edge of the second outer load path support structure of thesecond satellite, the second satellite stacked on top of the firstsatellite, and wherein the inner cavity of the first satellite isadapted to receive the booster rocket of the second satellite below anupper edge of the first outer load path support structure while thefirst outer load path support structure of the first satellitesubstantially contacts the second outer load path support structure ofthe second satellite.
 13. The stack of satellites of claim 12, whereinthe axial thrusters are positioned at a bottom of the first satellitebetween the primary axis and the outer cylinder, wherein the axialthrusters are oriented to impart impulses substantially parallel to theprimary axis, and wherein the axial thrusters are adapted to provide avelocity increment in a direction of the primary axis and to re-orientthe primary axis.
 14. The stack of satellites of claim 12, wherein theradial thrusters are positioned around the outer cylinder, wherein theradial thrusters are oriented to impart impulses that are substantiallyperpendicular to the primary axis, and wherein the radial thrusters areadapted to increase and decrease a rate of spin around the primary axisand to provide a velocity increment in a direction perpendicular to theprimary axis.
 15. The stack of satellites of claim 12, wherein thecanted thrusters are positioned around an outer surface of the firstsatellite, wherein the canted thrusters are oriented to impart impulsesthat substantially intersect a center of gravity of the first satellite,and wherein the canted thrusters are adapted to provide a velocityincrement in the direction of the canted thruster.
 16. The stack ofsatellites of claim 15, wherein a first set of the canted thrusters arepositioned above a plane that bisects the center of gravity and isperpendicular to the primary axis, and wherein a second set of thecanted thrusters are positioned below the plane.
 17. The stack ofsatellites of claim 16, wherein a cant angle between impulse vectors ofthe first set and the plane is in a range of about 30 degrees to about60 degrees.
 18. The stack of satellites of claim 16, wherein a cantangle between impulse vectors of the second set and the plane is in arange of about 30 degrees to about 60 degrees.
 19. The stack ofsatellites of claim 12, wherein the attitude control subsystem isfurther adapted to maintain the first satellite in an ecliptic normalattitude.
 20. The stack of satellites of claim 12, wherein the orbitmaintenance subsystem is further adapted to maintain the first satellitein a Molniya orbit with an orbital period of about 12 hours, an angle ofinclination of about 63.4 degrees, and a desired nodal separation of thefirst satellite, with respect to other satellites within a satellitefleet, of 360 degrees divided by a total number of satellites in thesatellite fleet.
 21. A stack of satellites comprising: a first satellitecomprising: a payload subsystem, which includes a downlink antennaadapted to produce an antenna pattern that is a figure of revolution; abus subsystem coupled to the payload subsystem, wherein the bussubsystem includes at least one booster rocket and one or more fueltanks; a first outer load path support structure coupled to the bussubsystem, wherein the first outer load path support structuresubstantially surrounds the payload subsystem and the bus subsystem,wherein the first outer load path support structure includes asubstantially cylindrical structure positioned at an outer radius of thefirst satellite; at least one booster rocket adapted to impart impulsessubstantially parallel to the primary axis; an attitude controlsubsystem adapted to maintain the first satellite in an ecliptic normalattitude during an operational life of the satellite; an orbitmaintenance subsystem to correct for orbital deviations during anoperational life of the first satellite, wherein the attitude controlsubsystem and the orbit maintenance subsystem includes a plurality ofaxial thrusters, radial thrusters, and canted thrusters; and a secondsatellite comprising a second outer load path support structurecomprising a single substantially cylindrical structure positioned at asingle outer radius of the second satellite; and wherein the first outerload path support structure is configured to couple to the second outerload path support structure of the second satellite and to substantiallybear a load of the second satellite, the second satellite being stackedabove the first satellite, a contact between the first outer load pathsupport structure of the first satellite and the second outer load pathsupport structure of the second satellite being substantially continuousalong the outer radius of the first outer load path support structureand the outer radius of the second outer load path support structure.